System for reducing thermal stresses in a leading edge of a high speed vehicle

ABSTRACT

A hypersonic aircraft includes one or more leading edge assemblies that are designed to manage thermal loads experienced at the leading edges during high speed or hypersonic operation. The leading edge assembly includes a plurality of structural layers and a plurality compliant layers alternately stacked with each other to facilitate thermal expansion and movement between the plurality of structural layers, while also providing a thermal break between the plurality of structural layers.

FIELD

The present subject matter relates generally leading edge technologiesfor use in high speed vehicles, such as hypersonic aircraft.

BACKGROUND

High speed vehicles often experience thermal management issues resultingfrom high heat load experienced during high speed operation,particularly at leading edges where the free air stream impinges on thevehicle. For example, in an application involving hypersonic aircrafts,the leading edges can include the nose, engine cowls, and the leadingedges of wings and stabilizers. Particularly when these vehicles areoperating in the hypersonic speed range (e.g., Mach 5 or greater), theleading edges may be subjected to very high heat load (e.g., 500-1500W/cm²) as the incident airflow passes through a bow shock and comes torest at the vehicle surface, converting the kinetic energy of the gas tointernal energy and greatly increasing its temperature. Unmitigatedexposure to such thermal loading can result in component degradationand/or failure.

Improvements in materials and manufacturing techniques have enabledhypersonic aircraft to operate at higher speeds and temperatures. Forexample, materials have been developed to increase the temperatures thata component can withstand while maintaining its structural integrity. Inthis regard, for example, nickel-based superalloys might be used to 800°C., single-crystal materials might be used to 1200° C., and refractorymetals may be required for even higher temperatures. In addition,various cooling technologies have been developed to provide cooling tothe leading edges of hypersonic vehicles. However, correspondingadvancements in vehicle speed and duration of high speed flight timeshave created the need for further improvement in the cooling ability andhigh temperature durability of the leading edges of high speed vehicles.

In addition, because the heat load is largest at the stagnation point orleading edge of a hypersonic vehicle, large temperature gradients maydevelop within the structure defining the leading edge. For example, theoutermost layer of leading edge may experience the highest heat andlargest thermal expansion, while the innermost layer may experiencerelatively low heat and less thermal expansion. As a result, thermalstresses may develop within the leading edge that can result incomponent failure or degradation.

Accordingly, improvements to hypersonic aircraft and propulsiontechnologies would be useful. More specifically, improvements in thermalmanagement technologies for reducing thermal stresses within the leadingedges of hypersonic vehicles would be particularly beneficial.

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one exemplary embodiment of the present disclosure, a leading edgeassembly for a hypersonic vehicle, the leading edge assembly comprising:an outer wall that is tapered to a forward end; a tip portion joined tothe forward end of the outer wall and extending forward toward a leadingedge, the tip portion comprising; an inner layer positioned at an aftend of the tip portion; an outer layer defining the leading edge at aforward end of the tip portion; and one or more compliant layerspositioned between the inner layer and the outer layer for facilitatingmovement between the inner layer and the outer layer.

According to another exemplary embodiment, a leading edge assembly for ahypersonic vehicle, the leading edge assembly comprising: an outer wallthat is tapered to a forward end; a tip portion joined to the forwardend of the outer wall, the tip portion comprising plurality ofstructural layers and a plurality compliant layers alternately stackedwith each other, wherein the plurality of compliant layers facilitatemovement between the plurality of structural layers.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures.

FIG. 1 is a close-up, cross-sectional, schematic view of a leading edgeof a hypersonic vehicle in accordance with an exemplary embodiment ofthe present disclosure.

FIG. 2 is a method of cooling a leading edge and facilitatingmagnetohydrodynamic generation or control of a hypersonic vehicleaccording to an exemplary embodiment of the present subject matter.

FIG. 3 is a close-up, cross-sectional, schematic view of a leading edgeof a hypersonic vehicle in accordance with another exemplary embodimentof the present disclosure.

FIG. 4 is a close-up, cross-sectional, schematic view of a leading edgeof a hypersonic vehicle in accordance with another exemplary embodimentof the present disclosure.

FIG. 5 is a close-up, cross-sectional, schematic view of a leading edgeof a hypersonic vehicle in accordance with another exemplary embodimentof the present disclosure.

FIG. 6 is a close-up, cross-sectional, schematic view of a leading edgeof a hypersonic vehicle in accordance with another exemplary embodimentof the present disclosure.

FIG. 7 is a close-up, cross-sectional, schematic view of an outer wallof a leading edge of a hypersonic vehicle in accordance with anotherexemplary embodiment of the present disclosure.

Repeat use of reference characters in the present specification anddrawings is intended to represent the same or analogous features orelements of the present invention.

DETAILED DESCRIPTION

Reference now will be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention.

The word “exemplary” is used herein to mean “serving as an example,instance, or illustration.” Any implementation described herein as“exemplary” is not necessarily to be construed as preferred oradvantageous over other implementations. Moreover, each example isprovided by way of explanation of the invention, not limitation of theinvention. In fact, it will be apparent to those skilled in the art thatvarious modifications and variations can be made in the presentinvention without departing from the scope of the invention. Forinstance, features illustrated or described as part of one embodimentcan be used with another embodiment to yield a still further embodiment.Thus, it is intended that the present invention covers suchmodifications and variations as come within the scope of the appendedclaims and their equivalents.

As used herein, the terms “first,” “second,” and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.The singular forms “a,” “an,” and “the” include plural references unlessthe context clearly dictates otherwise. The terms “coupled,” “fixed,”“attached to,” and the like refer to both direct coupling, fixing, orattaching, as well as indirect coupling, fixing, or attaching throughone or more intermediate components or features, unless otherwisespecified herein.

The terms “forward” and “aft” refer to relative positions within a gasturbine engine or vehicle, and refer to the normal operational attitudeof the gas turbine engine or vehicle. For example, with regard to a gasturbine engine, forward refers to a position closer to an engine inletand aft refers to a position closer to an engine nozzle or exhaust. Theterms “upstream” and “downstream” refer to the relative direction withrespect to fluid flow in a fluid pathway. For example, “upstream” refersto the direction from which the fluid flows, and “downstream” refers tothe direction to which the fluid flows.

Approximating language, as used herein throughout the specification andclaims, is applied to modify any quantitative representation that couldpermissibly vary without resulting in a change in the basic function towhich it is related. Accordingly, a value modified by a term or terms,such as “about,” “approximately,” and “substantially,” are not to belimited to the precise value specified. In at least some instances, theapproximating language may correspond to the precision of an instrumentfor measuring the value, or the precision of the methods or machines forconstructing or manufacturing the components and/or systems. Forexample, the approximating language may refer to being within a 10percent margin.

Here and throughout the specification and claims, range limitations arecombined and interchanged, such ranges are identified and include allthe sub-ranges contained therein unless context or language indicatesotherwise. For example, all ranges disclosed herein are inclusive of theendpoints, and the endpoints are independently combinable with eachother.

In general, aspects of the present subject matter are directed toleading edge assemblies for high speed aircraft or vehicles, such ashypersonic aircraft. As used herein, the term “hypersonic” refersgenerally to air speeds of about Mach 4 up to about Mach 10 or greater,such as Mach 5 and up. However, it should be appreciated that aspects ofthe present subject matter are not limited only to hypersonic flight,but may instead apply to applications involving other high speedvehicles, projectiles, objects, etc. The description of leading edgeassemblies herein with respect to use on a hypersonic aircraft are onlyexamples intended to facilitate the explanation of aspects of thepresent subject matter. The present subject matter is not limited tosuch exemplary embodiments and applications.

Notably, as explained above, high speed vehicles, such as a hypersonicaircraft, typically experience extremely high temperatures and thermalgradients during high speed or hypersonic operation. The temperaturegradients that are caused by the high heat flux are often a more severeproblem than the temperature itself. For example, the thermalconductivity of the structural material, in combination with the heatflux, sets the temperature gradient within the material, and at highheat loads this gradient leads to mechanical stresses that cause plasticdeformation or fracture of the material. The heat load to the structuralmaterial should be reduced to maintain the structural integrity of thecomponents.

As explained above, the leading edges of such high speed vehicles oftenexperience the highest thermal loading. For example, a hypersonicvehicle may include a plurality of leading edge assemblies (e.g.,identified generally herein by reference numeral 300) which experiencehigh thermal loads during hypersonic flight. In this regard, leadingedge assemblies 300 may be provided on a forward end of the aircraftwings, the nose cone, the vertical stabilizer, the engine cowls of thepropulsion engines, or other leading edges or surfaces of the hypersonicaircraft. According to exemplary embodiments of the present subjectmatter, leading edge assemblies 300 include features for mitigating theeffects of such thermal loading, e.g., by carrying heat out of theregion.

Notably, it is typically desirable to make leading edge assemblies 300as sharp or pointed as possible, e.g., in order to reduce drag on thehypersonic vehicle. However, referring now to FIG. 1, when leading edgeassemblies 300 are formed into a small radius of curvature, extremelyhigh temperatures and thermal gradients are experienced within leadingedge assembly 300 at its forward or leading edge, also referred toherein as a stagnation line, a stagnation point 302, or similar terms.In this regard, as a hypersonic vehicle is traveling through air athypersonic speeds, a free stream flow of air (e.g., identified herein byreference numeral 304) passes over and around leading edge assembly 300,thereby generating large thermal loads. Aspects of the present subjectmatter are directed to thermal management technologies and features forcooling leading edge assemblies 300, with a particular focus on theregions close to stagnation point 302, where the most serious thermalmanagement issues typically arise.

It should be appreciated that the leading edge assemblies 300illustrated herein are simplified cross section illustrations ofexemplary leading edges. The size, configuration, geometry, andapplication of such leading edge technologies may vary while remainingwithin the scope of the present subject matter. For example, the leadingedge assemblies 300 described herein define a radius of between about 1mm and 3 mm. However, according to alternative embodiments, leading edgeassemblies could have any other suitable radii.

The cooling technologies and thermal management features are describedherein as being used to cool portions of one or more parts of ahypersonic aircraft, such as the leading edges of the wings, nose,propulsion engines, or other parts of the hypersonic aircraft thatexperience large temperatures and thermal gradients. However, it shouldbe appreciated that aspects of the present subject matter may be used tomanage thermal loading such as high temperatures and thermal gradientswithin any component and in any suitable application. In this regard,for example, aspects of the present subject matter may apply to anyother hypersonic vehicle or to any other technology or system havingcomponents that are exposed to high temperatures and/or largetemperature gradients.

In addition, although various techniques, component configurations, andsystems are described herein for cooling leading edge assemblies 300 ofa hypersonic vehicle, it should be appreciated that variations andmodifications may be made to such technologies without departing fromthe scope of the present subject matter. In addition, one or more suchtechnologies may be used in combination with each other to achieveimproved cooling and thermal management. In this regard, although eachcooling technology is described in isolation in order to clearlydescribe how each technology functions, the embodiments described areonly examples intended for the purpose of illustration and explanation,and are not intended to limit the scope of the present subject matter inany manner.

In addition, according to exemplary embodiments of the present subjectmatter, some or all components described herein may be formed using anadditive-manufacturing process, such as a 3-D printing process. The useof such a process may allow certain components of a hypersonic vehicle,such as leading edge assemblies 300, to be formed integrally, as asingle monolithic component, or as any suitable number ofsub-components. As used herein, the terms “additively manufactured” or“additive manufacturing techniques or processes” refer generally tomanufacturing processes wherein successive layers of material(s) areprovided on each other to “build-up,” layer-by-layer, athree-dimensional component. The successive layers generally fusetogether to form a monolithic component which may have a variety ofintegral sub-components.

Although additive manufacturing technology is described herein asenabling fabrication of complex objects by building objectspoint-by-point, layer-by-layer, typically in a vertical direction, othermethods of fabrication are possible and within the scope of the presentsubject matter. For example, although the discussion herein refers tothe addition of material to form successive layers, one skilled in theart will appreciate that the methods and structures disclosed herein maybe practiced with any additive manufacturing technique or manufacturingtechnology. For example, embodiments of the present invention may uselayer-additive processes, layer-subtractive processes, or hybridprocesses.

Suitable additive manufacturing techniques in accordance with thepresent disclosure include, for example, Fused Deposition Modeling(FDM), Selective Laser Sintering (SLS), 3D printing such as by inkjets,laser jets, and binder jets, Sterolithography (SLA), Direct SelectiveLaser Sintering (DSLS), Electron Beam Sintering (EBS), Electron BeamMelting (EBM), Laser Engineered Net Shaping (LENS), Laser Net ShapeManufacturing (LNSM), Direct Metal Deposition (DMD), Digital LightProcessing (DLP), Direct Selective Laser Melting (DSLM), Selective LaserMelting (SLM), Direct Metal Laser Melting (DMLM), and other knownprocesses.

The additive manufacturing processes described herein may be used forforming components using any suitable material. For example, thematerial may be metal, concrete, ceramic, epoxy, or any other suitablematerial that may be in solid, liquid, powder, sheet material, wire, orany other suitable form or combinations thereof. More specifically,according to exemplary embodiments of the present subject matter, theadditively manufactured components described herein may be formed inpart, in whole, or in some combination of materials including but notlimited to pure metals, nickel alloys, chrome alloys, titanium, titaniumalloys, magnesium, magnesium alloys, aluminum, aluminum alloys, andnickel or cobalt based superalloys (e.g., those available under the nameInconel® available from Special Metals Corporation). These materials areexamples of materials suitable for use in the additive manufacturingprocesses described herein, and may be generally referred to as“additive materials.”

In addition, the additive manufacturing process disclosed herein allowsa single component to be formed from multiple materials. Thus, thecomponents described herein may be formed from any suitable mixtures ofthe above materials. For example, a component may include multiplelayers, segments, or parts that are formed using different materials,processes, and/or on different additive manufacturing machines. In thismanner, components may be constructed which have different materials andmaterial properties for meeting the demands of any particularapplication. In addition, although the components described herein areconstructed entirely by additive manufacturing processes, it should beappreciated that in alternate embodiments, all or a portion of thesecomponents may be formed via casting, machining, and/or any othersuitable manufacturing process. Indeed, any suitable combination ofmaterials and manufacturing methods may be used to form thesecomponents.

Referring still to FIG. 1, leading edge assembly 300 will be describedin more detail according to an exemplary embodiment of the presentsubject matter. Specifically, FIG. 1 provides a cross-sectional view ofa leading edge assembly 300, which may be positioned at a leading edge(e.g., a forward end, a leading end, upstream end, etc.) of anycomponent of a hypersonic aircraft. For example, leading edge assembly300 may be, e.g., a leading edge of an inlet duct to a hypersonicpropulsion engine, a leading edge of a turbine engine, a leading edge ofa wing of the aircraft, a nose of the aircraft, a forward end of avertical stabilizer, etc.

As explained herein, large thermal loads may be experienced by leadingedge assemblies 300 during hypersonic flight operations. As used herein,the terms “thermal load” and the like are intended generally to refer tothe high temperatures, temperature gradients, or heat flux experiencedwithin a component of a hypersonic or high-speed vehicle. According toexemplary embodiments of the present subject matter, leading edgeassemblies 300 are formed or provided with thermal regulation featuresor technologies for managing these thermal loads.

For example, as described in more detail below with reference to FIG. 1,leading edge assembly 300 may include features for providing a coolant,such as a cooling fluid or cooling material through an outer wall 320 ofleading edge assembly 300 to reduce a temperature of outer wall 320and/or a temperature gradient experienced within leading edge assembly300. FIGS. 2 through 5 provide cooling technologies for leading edgeassemblies 300 according to alternative embodiments. It should beappreciated that the thermal regulation features and technologiesdescribed herein for each exemplary leading edge assembly 300 may beused alone or in combination with any other leading edge technologiesdescribed herein to regulate the thermal loading on one or more leadingedge assemblies 300 of a hypersonic vehicle, or any other surface of anyother component that experiences high thermal loading.

As explained above, outer wall 320 and other components of leading edgeassembly 300 may be formed from any suitable material. According to anexemplary embodiment, such materials are selected to withstand the highthermal loading experienced by the leading edges of a hypersonicaircraft. For example, outer wall 320 may be constructed from at leastone of aluminum, titanium, titanium aluminide, tungsten, tungsten alloy,nickel superalloy, refractory metal, single-crystal metal, ceramic,ceramic matrix composite (CMC), ultra-high temperature ceramics (UHTCs,including high melting point diborides, nitrides, etc.), orcarbon-carbon composite. In addition, or alternatively, outer wall 320may include composites such as silicon carbide (SiC), SiC composites,carbon-fiber reinforced SiC matrix and other carbide matrix composites,composites with and without surface coatings, and/or high entropyalloys, including refractories, platinum group metals, hafnium alloys,etc. Nevertheless, it may still be desirable in certain applications toprovide additional cooling capacity for thermal management of the highheat loads experienced by leading edge assembly 300. Moreover, asexplained above, the additive manufacturing technologies may be used toprint leading edge assembly 300 (e.g. including outer wall 320) as asingle monolithic component, and may facilitate improved coolingtechnologies and leading edge features.

As is shown in the embodiment depicted, the outer wall 320 is generallyformed by a first wall section 322 and a second wall section 324. Morespecifically, the first wall section 322 and the second wall section 324each include outer surfaces together forming an outer surface 326 andinner surfaces together forming an inner surface 328. In addition, firstwall section 322 and second wall section 324 may be angled relative toeach other such that leading edge assembly 300 is tapered from an aftend 330 of leading edge assembly 300 to a forward end 332 of leadingedge assembly 300 (e.g., which corresponds to stagnation point 302). Inother words, leading edge assembly 300 is wider or taller proximate aftend 330 of leading edge assembly 300 and narrows as it approachesstagnation point 302. Notably, the taper angle may vary depending onaerodynamic and other considerations while remaining within the scope ofthe present subject matter. For example, according to exemplaryembodiments, leading edge assembly 300 may not be symmetric, e.g.,defining a sharper angle on one side.

As described above, for the embodiment shown, the outer surfaces 326 ofthe first wall section 322 and the second wall section 324 meet at astagnation point 302 and generally form a leading edge portion 340 ofthe outer wall 320. The leading edge portion 340 also defines at leastpart of outer surface 326 and inner surface 328. For the embodimentdepicted, the leading edge assembly 300 includes a cavity or a chamber342 positioned between the first wall section 322 and the second wallsection 324 and in fluid communication with the inner surface 328 of theleading edge portion 340. In addition, according to an exemplaryembodiment, leading edge portion 340 defines at least one fluidpassageway 350 that passes through outer wall 320.

More specifically, according to the illustrated embodiment, flowpassageway 350 is a hole or aperture that is defined through outer wall320 such that it extends from an inlet 352 positioned at inner surface328 to and outlet 354 positioned at outer surface 326. In addition, asillustrated, outlet 354 is positioned at stagnation point 302 or atforward end 332 of leading edge assembly 300. Although a single fluidpassageway 350 is illustrated in the cross-section depicted in FIG. 1,it should be appreciated that fluid passageway 350 may instead be anelongated slot extending along a leading edge or stagnation line ofleading edge assembly 300. According still other embodiments, leadingedge portion 340 of leading edge assembly 300 may include a plurality offluid passageways 350 spaced apart from each other along a length ofleading edge assembly 300, e.g., such as along a span of aircraft wings.

Indeed, fluid passageways 350 may be defined through outer wall 320 inany suitable number, geometry, configuration, etc. For example, leadingedge assembly 300 may generally define a radial direction R and acircumferential direction C. The radial direction R that extends outwardfrom a center of curvature (not labeled) of leading edge portion 340within the cross-sectional plane illustrated in FIG. 1. Thecircumferential direction C generally wraps around leading edge assembly300 from first wall section 322 to second wall section 324 as shown inFIG. 1. In addition, leading edge assembly 300 may define a longitudinaldirection L, which according to the exemplary embodiment describedherein, corresponds to the longitudinal direction L of a hypersonicaircraft and is substantially parallel to an angle of attack definedbetween leading edge assembly 300 and the primary direction of the flowof air 304.

According to the illustrated embodiment, fluid passageway 350 extendsthrough outer wall 320 along the radial direction R. However, it shouldbe appreciated that according to alternative embodiments, fluidpassageways 350 may pass through outer wall 320 at an angle other thanthe radial direction R, may pass along a length of first wall section322 and/or second wall section 324, may terminate at other locationsboth on inner surface 328 and outer surface 326, may vary incross-sectional size or profile, or may vary in any other suitablemanner for providing a flow of cooling fluid 360 to a desired locationwithin or on leading edge assembly 300.

In this regard, leading edge assembly 300 may further include a coolantsupply 362 for providing the flow of cooling fluid 360 through fluidpassageway 350 and out of leading edge assembly 300. As used herein, theterm “cooling fluid,” “coolant,” and the like are generally intended torefer to any material that passes through outer wall 320 to cool outersurface 326 of leading edge assembly 300. According to an embodiment ofthe present subject matter, as described in more detail below, coolingfluid 360 may include at least one of air, water, steam, liquid metal,carbon dioxide, argon, helium, etc.

It should be further appreciated that the cooling fluid 360 is notlimited to a particular state or phase of material. In this regard,although the terms “liquid” and “fluid” may be used to refer to coolingfluid 360, cooling fluid 360 could also include solid materials, vapors,etc. Furthermore, the present subject matter is not limited by thesource of cooling fluid 360. For example, cooling fluid 360 may bepumped from a storage reservoir, may be contained within leading edgeassembly 300 as a solid material that liquefies at a criticaltemperature, seeps through leading edge portion 340 or fluid passageway,vaporizes to absorb thermal energy, and carries away thermal energy asit flows downstream.

According to an exemplary embodiment of the present subject matter,cooling fluid 360 may be solid or liquid metals. For example, accordingto an exemplary embodiment, cooling fluid 360 is an alkali metal.Exemplary liquid metals that may be pumped through leading edge assembly300 as cooling fluid 360 may include at least one of cesium (Cs, with aboiling point of approximately 678° C.), potassium (K, with a boilingpoint of approximately 774° C.), sodium (Na, with a boiling point ofapproximately 883° C.), lithium (Li, with a boiling point ofapproximately 1347° C.), indium (In, with a boiling point ofapproximately 2000° C.), and/or gallium (Ga, with a boiling point ofapproximately 2205° C.). In addition, or alternatively, cooling fluid360 may include ammonia (NH₃) or sulfur dioxide (SO₂). According toexemplary embodiments, cooling fluid 360 may be a mixture or combinationof such liquid metals or compositions.

In addition, cooling fluid 360 may be selected based on its materialproperties, such as an ionization potential. The ionization potential ofcooling fluid 360 may be important to facilitate magnetohydrodynamiccontrol or generation using leading edge assembly 300 (as explainedbelow). For example, according to an exemplary embodiment, the liquidmetal cooling fluid 360 may have an ionization potential of betweenabout 1 and 10 electronvolts (eV), between about 2 and 8 electronvolts(eV), between about 4 and 6 electronvolts (eV), or about 5 electronvolts(eV). According to exemplary embodiments, cooling fluid 360 is selectedsuch that its ionization potential is much lower than the ionizationpotential for air. For example, the ionization potential for O₂ isapproximately 12.6 eV and the ionization potential for N₂ isapproximately 14.5 eV.

These materials can function as ‘seeds’ that ionize at much lowertemperature than air. It should be appreciated that the term“ionization” and the like are used generally herein to refer to bothcomplete and partial ionization. For example, the term ionization mayrefer to a partial or weak ionization, e.g., where the density of ionsis a small fraction (say, 10⁻⁵) of the density of the neutral species.In addition, it should be appreciated that it is not only the seedspecies that can ionize, but instead that any present species mayionize, including oxygen and nitrogen, if the temperature issufficiently high (e.g., due to a high Mach number).

According to alternative embodiments, cooling fluid 360 may be selectedbased on another material property, such as boiling temperature.According to an exemplary embodiment, cooling fluid 360 may have aboiling temperature that is less than a melting temperature of outerwall 320. In this manner, outer wall 320 or other portions of leadingedge assembly 300 may be impregnated with solid materials (e.g., solidmetals) that passively melt and flow as a cooling fluid 360. This meltedmetal cooling fluid 360 will vaporize and begin cooling outer wall 320before outer wall 320 begins melting. In this manner, the vaporizationof cooling fluid 360 on an outer surface 326 may maintain thetemperature of outer wall 320 at the boiling point or boilingtemperature of cooling fluid 360. If this boiling point is selected tobe below a melting temperature of outer wall 320, the structuralintegrity of outer wall 320 may be maintained until the cooling fluid360 has been used up. For example, according to one embodiment, theliquid metal cooling fluid 360 has a boiling temperature of greater thanabout 500° C., greater than about 700° C., greater than about 900° C.,greater than about 1100° C., or about 750° C. In addition, oralternatively, the boiling temperature of the liquid metal cooling fluid360 may be below about 2000° C., below about 1500° C., below about 1000°C., or below about 800° C. For example, according to an exemplaryembodiment, cooling fluid 360 is lithium (Li), which boils at about1342° C.

As used herein, the term “boiling point,” “boiling temperature,” and thelike are referring to temperatures tabulated at 1 atm ambient pressure.However, it should be appreciated that such boiling temperatures may bedifferent at flight conditions, e.g., when ambient and experiencedpressures are different. In addition, although the boiling point ofcooling fluid 360 is described as being below the melting temperature ofouter wall 320, it should be appreciated that it may be desirable tohave the boiling point of cooling fluid 360 even lower, e.g.,corresponding to some predetermined structure temperature where theintegrity of outer wall 320 begins to degrade or weaken.

It should be appreciated that the cooling fluids 360 described hereinare only exemplary and are not intended to limit the scope of thepresent subject matter. For example, cooling fluids 360 may be selectedto be liquid or solid, may have varying boiling points and ionizationpotentials, and may exhibit other material properties as well. Forexample, it may be desirable to having materials with a relatively lowmass density, particularly for applications involving an aircraft. Inaddition, the heat of vaporization may be important to determine volumeand mass of required cooling fluid 360, e.g., lithium (Li) has auniquely high specific heat of vaporization (per unit mass). Accordingto exemplary embodiments, cooling fluids 360 may be selected to melt atreasonably low melting temperature, e.g., to minimize the complexityrequired of an onboard material handling system.

In addition, or alternatively, cooling fluid 360 may be a metal phasechange material. For example, the coolant may be a metal configured tochange from a solid phase to liquid or gas phase when exposed totemperatures generated during operation of a hypersonic propulsionengine during hypersonic flight operations. Additionally, oralternatively, the coolant may be a metal configured to change from aliquid phase to a gas phase when exposed to temperatures generatedduring operation of the hypersonic propulsion engine during hypersonicflight operations. However, in other embodiments other suitable coolantmay be utilized.

As best illustrated in FIG. 1, coolant supply 362 is in fluidcommunication with fluid passageway 350 for selectively providing a flowof coolant 360 through fluid passageway 350. Specifically, according tothe illustrated embodiment, coolant supply 362 includes a coolantreservoir 364 and a mechanical pump 366 is configured for pressurizingor supplying cooling fluid 360 to fluid passageway 350. According tostill other embodiments, cooling fluid 360 may be stored within apressurized tank for delivery to fluid passageway 350, or leading edgeassembly 300 may include a system of capillary tubes or anothercapillary structure for driving cooling fluid 360. Alternatively, thecooling fluid 360 may be transported from any suitable location in anyother suitable manner. For example, as illustrated, coolant supply 362generally provides a flow of pressurized cooling fluid 360 into chamber342 which is in fluid communication with fluid passageway 350. However,it should be appreciated that according to alternative embodiments,coolant supply 362 may include one or more conduits that are directlyfluidly coupled to fluid passageway 350. Moreover, although coolingfluid 360 is illustrated as a single flow from a single coolantreservoir 364, it should be appreciated that one or more coolingfluid(s) 360 may be provided from a plurality of coolant reservoirs 364,may include different cooling fluid(s) 360 at different pressures, etc.

As explained above, according to an exemplary embodiment of the presentsubject matter, the flow of cooling fluid 360 may include liquid metal.Notably, when the liquid metal cooling fluid 360 is exposed to such hightemperatures at leading edge assembly 300, it may ionize after exitingfluid passageway 350 to generate a flow of ionized vapor (e.g., asidentified generally by reference numeral 370 in FIG. 1). This flow ofionized the vapor 370 passes from stagnation point 302 downstream onleading edge assembly 300. According to an exemplary embodiment, thisflow of ionized vapor 370 may facilitate a magnetohydrodynamic processfor power generation and/or vehicle control. Specifically, according toan exemplary embodiment, leading edge assembly 300 may include amagnetohydrodynamic generator 372 (“MHD generator”) that interacts withthe flow of ionized vapor 370 to generate electricity or thrust forvehicle control.

In general, MHD generator 372 may be any device suitable for interactingwith the flow of ionized vapor 370 to generate electricity or thrust forvehicle control. It will be appreciated that a MHD generator 372 maygenerally utilize a magnetohydrodynamic converter that utilizes aBrayton cycle to transform thermal energy and kinetic energy directlyinto electricity. For example, MHD generator 372 may include one or moremagnet coils and/or electrodes positioned proximate to or within outerwall 320. According to an exemplary embodiment, the MHD generator 372may use the coils to generate electricity as the flow of ionized vapor360 passes through a magnetic field generated by the coil. In addition,or alternatively, MHD generator 372 may energize the magnet coils orelectrodes to interact with the flow of ionized vapor 370 in a mannerthat controls or adjusts flight by providing thrust to outer wall 320.MHD generator 372 may further include a controller and unique structurefor implementing control methodologies for steering during high speedflight, such as during hypersonic flight.

Specifically, as best illustrated in FIG. 2, a method of operatingleading edge assembly 300 may include a magnetohydrodynamic controlmethod. Specifically, method 400 includes, at step 410, providing a flowof coolant through at least one fluid passageway defined through aleading edge of a hypersonic vehicle, wherein the flow of coolantcomprises liquid metal (or any other suitable seed material) that isionized after exiting the at least one fluid passageway and passesdownstream as a flow of ionized vapor. Once the flow of ionized vapor isachieved, step 420 includes generating electricity or thrust for vehiclecontrol using a magnetohydrodynamic generator that interacts with theflow of ionized vapor.

In addition, or alternatively, as best shown in FIGS. 3 and 4, theleading edge portion 340 may be configured as a porous leading edgeportion, identified herein as porous tip 500. According to an exemplaryembodiment, the leading edge assembly 300 is configured to provide aflow of cooling fluid 360 through the chamber 342 to the inner surface328 of the leading edge portion 340, such that the flow of cooling fluid360 may seep through the porous tip 500 and cool the leading edgeportion 340 during operation of a hypersonic vehicle or a hypersonicpropulsion engine, e.g., during hypersonic flight operations. In thisregard, fluid passageways 350 may be defined through the porousstructure that composes or defines porous tip 500. In describing figuresherein, it should be appreciated that like reference numerals may beused to refer to the same or similar features between embodiments.

Specifically, as shown in FIGS. 3 and 4, an outer wall 320 of leadingedge assembly 300 includes a first wall section 322 and a second wallsection 324, each of which extend from the a forward end 502 and an aftend (which may correspond, for example, with aft end 330 of leading edgeassembly 300) along the longitudinal direction L. Notably, forward ends502 of first wall section 322 and second wall section 324 may stop shortof stagnation point 302. In addition, porous tip 500 may be joined toforward ends 502 of first wall section 322 and second wall section 324.In this regard, porous tip 500 may generally be defined as a curvedregion coupled to first wall section 322 and second wall second 324,which are illustrated as being substantially straight.

According to exemplary embodiments, for example, first wall section 322,second wall section 324, and tip portion 500 may be simultaneouslyadditively manufactured as a single, integral, and monolithic piece.However, as will be described in more detail below, outer wall 320 maybe substantially solid and hermetic, while porous tip 500 may include aporous structure for permitting a flow of cooling fluid 360 to passtherethrough toward forward end 332 of leading edge assembly 300.

According to an exemplary embodiment, the leading edge portion 340, ormore specifically porous tip 500, may define a constant porosity forpassing cooling fluid 360. As used herein, the term “porosity” may beused generally to refer to a measure of the void or empty spaces withina material or structure. Thus, a structure having porosity has openpassages, cells, or structures through which fluidly may flow from oneporous cell to another. For example, porosity may be used to refer to afraction of the volume of voids or open space over a total volume of acomponent. According to exemplary embodiments, the porosity of poroustip 500 may be greater than about 5%, 10%, 20%, 40% or greater than even50%. In addition, or alternatively, the porosity of porous tip 500 maybe less than about 80% 60%, 40%, 20%, or 5%. It should be appreciatedthat the porosity of porous tip 500 may vary depending on theapplication while remaining within scope of the present subject matter.For example, the porosity may vary based on the mass flow rate of thecooling fluid 360, the mechanical properties of porous tip 500, based onanticipated flight conditions, or based on any other suitableparameters.

Notably, according to the illustrated embodiment, porous tip may definea variable porosity, e.g., in order to concentrate a cooling fluid 360proximate the stagnation point 302. More specifically, the porous tip500 may include a first porous region 510, a second porous region 512,and a third porous region 514. According to the illustrated embodiment,first porous region 510 is positioned at forward end 322 of leading edgeassembly 300 and includes the stagnation point 302. By contrast, each ofsecond porous region 512 and third porous region 514 are positioneddownstream of first porous region 510 or stagnation point 302. Accordingto an exemplary embodiment, a first porosity of first porous region 510is different than a second porosity of second porous region 512. Morespecifically, according to an exemplary embodiment, the first porosityis greater than the second porosity. Similarly, the third porosity maybe the same as or different and the second porosity, but is also lessthan the first porosity. For example, according to an exemplaryembodiment, the first porosity may be at least about 10 percent greaterthan the second and third porosity, such as at least about 25 percentgreater, such as at least about 50 percent greater, such as at leastabout 100 percent greater, and up to about 1000 percent greater porositythan the second and third porosity.

Similar to the embodiments described above, coolant supply 362 may be influid communication with chamber 342 and/or porous tip 500 forselectively providing a flow of coolant 360 through porous tip 500.Specifically, according to the illustrated embodiment, inner surface 328of porous tip 500 is exposed to a pressurized cooling fluid 360 withinchamber 342. However, the amount of cooling fluid 360 that flows throughfirst porous region 510 may be greater than the amount of cooling fluid360 flowing through one or both of second porous region 512 and thirdporous region 514. Notably, the ratio or amount of cooling fluid 360flowing through each region of porous tip 500 may be adjusted bymanipulating the porosities within each region 510-514. According to theillustrated embodiment, it is desirable to have the highest porosity infirst porous region 510 to direct the largest amount of cooling fluid360 toward the most heat affected region, i.e., stagnation point 302. Inaddition, higher porosity at the stagnation point 302 may helpcompensate for the fact that the pressure is highest at this point,which decreases the amount of cooling fluid 360 that naturally flowstoward stagnation point 302.

Notably, porous tip 500 is illustrated as having three distinct porousregions 510-514. However, it should be appreciated that using theadditive manufacturing techniques described herein, porous tip 500 mayhave a progressively varying porosity, i.e., such that the porositycontinuously and progressively increases from forward end 502 of outerwall 320 to the highest porosity at stagnation point 302. In thismanner, porous tip 500 may be envisioned as having 10, 20, 50, 100, oreven more subregions, each of which has a progressively increasingporosity as they approach stagnation point 302. As illustrated, each ofthese subregions extends from the inner surface 328 to outer surface 326along the radial direction R and has a substantially constant porosityalong the radial direction R. In addition, according to alternativeembodiments, one or more of these regions, such as porous regions510-514, may also vary in porosity along the radial direction R.

According to exemplary embodiments, porous tip 500 may includeadditional features for directing the flow of cooling fluid 360 to thedesired locations within leading edge assembly 300. In this regard,referring for example to FIG. 4, porous tip 500 may include a pluralityof internal barriers 520 that separate one or more of porous regions510-514. In this regard, as illustrated, internal barriers 520 arestraight and solid walls that extend substantially along the radialdirection R from inner surface 328 to outer surface 326. However,according to alternative embodiments, porous tip 500 may have moreinternal barriers 520, the internal barriers 520 may extend in otherdirections through outer wall 320, may have varying thicknesses, etc.Other configurations are possible and within the scope of the presentsubject matter.

Referring now specifically to FIG. 4, leading edge assembly 300 mayinclude a nose cover 530 that is positioned at least partially overporous tip 500 to restrict the flow of cooling fluid 360 from coolantreservoir 364 and/or chamber 342. Specifically, nose cover 530 preventscooling fluid 360 from escaping porous tip 500 or leading edge portion340, but may be made from a material that ablates or melts away when theleading edge assembly 300 is exposed to a predetermined criticaltemperature. In this regard, for example, nose cover 530 may beconstructed from a material that melts before porous tip 500 and outerwall 320 reach their melting temperatures. Notably, when nose cover 530melts, cooling fluid 360 is released to cool or maintain the temperatureof leading edge assembly 300 at or below the melting point of porous tip500 and/or outer wall 320. Thus, according to an exemplary embodiment,the predetermined critical temperature at which nose cover 530 melts isbelow a temperature at which the structural integrity of leading edgeassembly 300 begins to deteriorate. Although nose cover 530 isillustrated as being used with a porous tip 500 having variableporosity, it should be appreciated that these two features may be usedtogether or independently of each other.

According to the illustrated embodiment, the nose cover 530 extends overthe entire porous tip 500, e.g., from forward end 502 of first wallsection 322 along the circumferential direction C to forward end 502 ofsecond wall section 324. Notably, according to an exemplary embodiment,nose cover 530 is impermeable by the flow of cooling fluid 360.Therefore, cooling fluid 360 is contained within porous tip 500, coolantreservoir 364, and/or chamber 342 until the nose cover 530 melts, whichcorresponds to a time when cooling of leading edge assembly 300 isneeded. According to alternative embodiments, nose cover 530 may coverless than the entire portion of porous tip 500. For example, nose cover530 may cover only a forward end 332 of first porous region 510.Alternatively, nose cover 530 may cover only second porous region 512and third porous region 514. According to still other embodiments, athickness of nose cover 530 (which may be measured substantially alongthe radial direction R), may vary depending on the circumferentiallocation along porous tip 500. According to still other embodiments,nose cover 530 may be a plug positioned at outlet 354 of fluidpassageway 350 (e.g., as shown in FIG. 1). In this regard, for example,nose cover 530 may extend at least partially within fluid passageway350. Other variations and modifications to nose cover 530 may be madewhile remaining within the scope of the present subject matter. Forexample, it should be appreciated that nose cover 530 can be made byfilling pores with a different material to block the pores and form nosecover 530, by attaching a separate cap element to leading edge assembly300, etc.

Referring now to FIG. 5, leading edge assembly 300 may include an aftbulkhead 532 positioned at an aft end 330 of outer wall 320 andextending substantially perpendicular to a longitudinal direction L. Asshown, chamber 342 is enclosed and defined between the aft bulkhead 532,first wall section 322, second wall section 324, and nose cover 530.Thus, chamber 342 may be a constant volume chamber or reservoir. Forexample, aft bulkhead 532 may connect to the rest of a hypersonicvehicle, may be a replaceable component, etc. According to such anexemplary embodiment, chamber 342 may be charged with cooling fluid 360.Notably, as leading edge assembly 300 heats up during hypersonicoperation, nose cover 530 may slowly melt while the pressure withinchamber 342 increases due to the increased temperature of the coolingfluid 360. Once the temperature of leading edge assembly 300 has reachedthe critical temperature, nose cover 530 melts away and chamber 342 issufficiently pressurized to drive the flow of cooling fluid 360 throughporous tip 500 to cool the leading edge assembly 300.

According to still other embodiments, porous tip 500 may be filled witha material that seeps out of porous tip 500 when nose cover 530 melts orablates away. More specifically, according to an exemplary embodiment,porous tip 500 may be filled with a metal cooling fluid 360 (e.g., insolid form) that may have a relatively low melting point, such that themetal filling the pores of porous tip 500 is configured to melt duringoperation of a hypersonic propulsion engine or a hypersonic aircraftduring high temperature operations, such as hypersonic flightoperations. Once the metal cooling fluid 360 that fills the pores ofporous tip 500 is melted, the cooling fluid 360 may flow through poroustip 500 in a similar manner as described above with reference to FIG. 4.

Referring now specifically to FIGS. 6 and 7, leading edge assembly 300will be described according to another exemplary embodiment of thepresent subject matter. As described below, leading edge assembly 300may be formed from one or more walls that include a layered,multi-functional material with compliant interfaces that are used toalleviate extreme thermal stresses experienced by a hypersonic aircraftat high heat flux locations, e.g., proximate stagnation point 302. Aswill be described below, each of the plurality of layers within thelayered wall may be made from the same material (e.g., such as metallicor ceramic materials) or may be made up of different materials.Moreover, each layer may be tailored to meet specific needs for a givenapplication, e.g., based on the expected thermal loading, theanticipated temperature gradient, the presence of localized coolingfeatures, etc.

As explained above, leading edge assembly 300 may include variouscooling technologies, such as technologies to provide transpirationcooling features that may include a cooling fluid 360 positioned withina cavity 342. Notably, in such an embodiment, the largest temperaturesmay be experienced on outer surface 326 proximate stagnation point 302and the lowest temperatures may be experienced proximate inner surface328, which may be exposed to cooling fluid 360. Thus, the largetemperature gradient experienced across outer wall 320 may result insignificant thermal stresses, e.g., due to the varied thermal expansionbetween the layers or regions of outer wall 320.

As shown in FIG. 6, leading edge assembly 300 may include a tip portion550 (e.g., which may correspond to leading edge portion 340) that isjoined to a forward end 502 of outer wall 320 or otherwise is positionedproximate forward end 332 leading edge assembly 300. As shown, tipportion 550 includes a plurality of structural layers 552 and aplurality of compliant layers 554 that are alternately stacked with eachother. In this manner, the plurality of compliant layers 554 mayfacilitate some movement between the plurality of structural layers 552and may also provide a thermal break or insulating gap between theplurality of structural layers 552. According to the illustratedembodiment, the plurality of compliant layers 554 are embedded withintip portion 550 between the plurality of structural layers 552 and arespaced apart along a thickness of the outer wall 320 or tip portion 550.

Although layers 552, 554 are described herein as being “structural” and“compliant,” it should be appreciated that these terms are used only todistinguish the various layers and are not intended to limit such layersto a particular material or material property. In addition, it should beappreciated that the term “layers” of outer wall 320 is used herein torefer generally to different regions within outer wall 320. In thisregard, each layer may be substantially isothermal or quasi-isothermal,e.g., such that each layer experiences substantially similar thermalloading. However, the term “layers” is not intended to limit the mannerof constructing leading edge assembly 300, or to otherwise indicate orrequire a precisely layered or laminar construction. For example,“layers” of leading edge assembly 300 may all be additively manufacturedas a single, integral piece using one or more materials having differentdensities, porosities, coefficients of thermal expansion, or othervarying material properties.

The plurality of structural layers 552 and the plurality of compliantlayers 554 may be formed from any suitable materials. For example,layers 552, 554 may be made from a metal, a ceramic, a ceramic matrixcomposite material, or any other suitable materials described herein. Inaddition, according to an exemplary embodiment, the plurality ofstructural layers 552 may be formed from the same material and may havethe same material properties. By contrast, according to alternativeembodiments, the plurality of structural layers 552 may have differentmaterials, compositions, or constructions, for best managing thermalloading experienced by leading edge assembly 300. Similarly, theplurality of compliant layers 554 may each be formed from the samematerial or from different materials which may or may not includematerials used to form the plurality of structural layers 552.

For example, referring now specifically to FIG. 7, a close up view oftip portion 550 will be described according to an exemplary embodimentof the present subject matter. As shown, tip portion 550 includes aninner layer 560 (i.e., one of a plurality of structural layers 552)positioned proximate an aft end of tip portion 550. In this regard,inner layer 560 may at least partially define inner surface 328 andchamber 342 of leading edge assembly 300. In addition, tip portion 550includes an outer layer 562 that defines the leading edge or stagnationpoint 302 at a forward end 332 of leading edge assembly 300.

Between inner layer 560 and outer layer 562, tip portion 550 may includeone or more compliant layers 554. For example, compliant layers 554 mayfacilitate movement between inner layer 560 and outer layer 562. Morespecifically, as shown in FIG. 7, tip portion 550 further includes anintermediate layer 564 and two compliant layers 554. More specifically,a first compliant layer 566 (i.e., one of a plurality of compliantlayers 554) is positioned between inner layer 560 and intermediate layer564. In addition a second compliant layer 568 is positioned betweenintermediate layer 564 and outer layer 562. It should be appreciatedthat the tip portion 550 may include additional structural layers 552and/or compliant layers 554.

It should be appreciated that the layered structure used to define tipportion 550 may be used in other regions of leading edge assembly 300,elsewhere within a hypersonic aircraft, or in other applications wherehigh thermal loading is expected. In addition, although structurallayers 552 and compliant layers 554 are illustrated as being curvedbetween first wall section 322 and second wall section 324, these layers552, 554 may take any other suitable shape, size, or position forfacilitating improved management of thermal stresses.

According to an exemplary embodiment, one or more of the structurallayers 552 and one or more of the compliant layers 554 may havedifferent material densities. In this manner, adjusting the materialdensity may affect the stresses induced within a material layer, mayadjust the thermal conductivity of the layer, etc. In addition, one ormore of the structural layers 552 and one or more of the compliantlayers 554 may have different coefficients of thermal expansion. In thismanner, it may be desirable to form outer layer 562 to have a lowercoefficient of thermal expansion than inner layer 560, e.g., to reducethe expansion difference experienced between these two layers whenexposed to different temperatures during hypersonic operation.

According to exemplary embodiments, each of the plurality of compliantlayers 554 may define a thickness 570. According to certain embodiments,the thickness 570 may be less than about 1 millimeter. The plurality ofcompliant layers 554 between the plurality of structural layers 552 mayeffectively act to distribute heat at, e.g., the stagnation point 302along the outer surface 326 to reduce a concentration of the heat at thestagnation point 302. The plurality of compliant layers 554 may be acavity with internal volume defined by the thickness 570 and may befilled with a fluid or material layer which has relatively high heattransfer coefficient, such as liquid sodium. Notably, in at leastcertain exemplary embodiments, the plurality of compliant layers 554 maydefine a smaller thickness 570, and a thickness of the material betweenthe plurality of compliant layers 554 may be less than or equal to about1 millimeter.

It should further be appreciated that the layered tip portion describedwith respect to FIGS. 6 and 7 may be used along with other technologiesdescribed herein to minimize the effects of thermal loading of leadingedge assembly 300. For example, according to an exemplary embodiment,tip portion 550 may define at least one cooling passageway (e.g., suchas fluid passageway 350 from FIG. 1) for providing a flow of coolingfluid 360 to the outer surface 326 of leading edge assembly 300. Leadingedge assembly 300 may further include a coolant supply, such as coolantsupply 362, for providing the flow of cooling fluid 360 to tip portion550. Furthermore, tip portion 550 may be divided into subregions,similar to porous tip 500, with each region having a different number,type, material, and configuration of structural layers 552 and compliantlayers 554. Other variations and modifications may be made whileremaining within the scope of the present subject matter.

Further aspects of the invention are provided by the subject matter ofthe following clauses:

1. A leading edge assembly for a hypersonic vehicle, the leading edgeassembly comprising: an outer wall that is tapered to a forward end; atip portion joined to the forward end of the outer wall and extendingforward toward a leading edge, the tip portion comprising; an innerlayer positioned at an aft end of the tip portion; an outer layerdefining the leading edge at a forward end of the tip portion; and oneor more compliant layers positioned between the inner layer and theouter layer for facilitating movement between the inner layer and theouter layer.

2. The leading edge assembly of any preceding clause, wherein the tipportion further comprises: an intermediate layer positioned between theinner layer in the outer layer, and wherein the one or more compliantlayers comprises a first compliant layer positioned between the innerlayer and the intermediate layer and a second compliant layer positionedbetween the intermediate layer and the outer layer.

3. The leading edge assembly of any preceding clause, wherein the outerwall includes a first wall section and a second wall section separatedby a chamber, and wherein the inner layer, the intermediate layer, andthe outer layer are each curved and extend between the first wallsection and the second wall section.

4. The leading edge assembly of any preceding clause, wherein the innerlayer and the outer layer are made from different materials.

5. The leading edge assembly of any preceding clause, wherein the outerlayer and the inner layer are made from a metal, a ceramic material, ora ceramic matrix composite material.

6. The leading edge assembly of any preceding clause, wherein the outerlayer and the inner layer have different densities.

7. The leading edge assembly of any preceding clause, wherein at leasttwo of the one or more compliant layers are made from differentmaterials.

8. The leading edge assembly of any preceding clause, wherein the innerlayer and the outer layer have a first coefficient of thermal expansionand the one or more of compliant layers have a second coefficient ofthermal expansion, the second coefficient of thermal expansion beinggreater than the first coefficient of thermal expansion.

9. The leading edge assembly of any preceding clause, wherein at leastone cooling passageway is defined through the inner layer, the outerlayer, and the one or more compliant layers.

10. The leading edge assembly of any preceding clause, furthercomprising: a coolant supply in fluid communication with the at leastone cooling passageway for selectively providing a flow of coolantthrough the at least one cooling passageway.

11. The leading edge assembly of any preceding clause, wherein theleading edge assembly is positioned on a wing, a nosecone, an enginecowl, an engine inlet, a fuselage, or a stabilizer of the hypersonicvehicle.

12. A leading edge assembly for a hypersonic vehicle, the leading edgeassembly comprising: an outer wall that is tapered to a forward end; atip portion joined to the forward end of the outer wall, the tip portioncomprising plurality of structural layers and a plurality compliantlayers alternately stacked with each other, wherein the plurality ofcompliant layers facilitate movement between the plurality of structurallayers.

13. The leading edge assembly of any preceding clause, wherein the outerwall includes a first wall section and a second wall section separatedby a chamber, and wherein the plurality of structural layers are curvedand extend between the first wall section and the second wall section.

14. The leading edge assembly of any preceding clause, wherein theplurality of structural layers are made from different materials.

15. The leading edge assembly of any preceding clause, wherein theplurality of structural layers are made from a metal, a ceramicmaterial, or a ceramic matrix composite material.

16. The leading edge assembly of any preceding clause, wherein at leasttwo of the plurality of structural layers have different densities.

17. The leading edge assembly of any preceding clause, wherein at leasttwo of the plurality of compliant layers are made from differentmaterials.

18. The leading edge assembly of any preceding clause, wherein theplurality of structural layers each have a first coefficient of thermalexpansion and the plurality of compliant layers each have a secondcoefficient of thermal expansion, the second coefficient of thermalexpansion being greater than the first coefficient of thermal expansion.

19. The leading edge assembly of any preceding clause, wherein at leastone cooling passageway is defined through the plurality of structurallayers and the plurality of compliant layers.

20. The leading edge assembly of any preceding clause, furthercomprising: a coolant supply in fluid communication with the at leastone cooling passageway for selectively providing a flow of coolantthrough the at least one cooling passageway.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal language of the claims.

What is claimed is:
 1. A leading edge assembly for a hypersonic vehicle,the leading edge assembly comprising: an outer wall that is tapered at aforward end thereof; and a tip portion joined to the forward end of theouter wall and extending forward toward a leading edge, the tip portioncomprising: a forward end that is curved, an aft end that is curved, aninner layer positioned at the aft end of the tip portion and conformingto the aft end of the tip portion, an outer layer defining the leadingedge at a forward end of the tip portion and conforming to the forwardend of the tip portion, and one or more compliant layers that is curvedand positioned between the inner layer and the outer layer forfacilitating movement between the inner layer and the outer layer, andwherein the inner layer and the outer layer have a first coefficient ofthermal expansion and the one or more of compliant layers have a secondcoefficient of thermal expansion, the second coefficient of thermalexpansion being greater than the first coefficient of thermal expansion.2. The leading edge assembly of claim 1, wherein the tip portion furthercomprises: an intermediate layer positioned between the inner layer andthe outer layer, and wherein the one or more compliant layers comprisesa first compliant layer positioned between the inner layer and theintermediate layer and a second compliant layer positioned between theintermediate layer and the outer layer.
 3. The leading edge assembly ofclaim 2, wherein the outer wall includes a first wall section and asecond wall section separated by a chamber, and wherein the inner layer,the intermediate layer, and the outer layer are each curved and extendbetween the first wall section and the second wall section.
 4. Theleading edge assembly of claim 1, wherein the inner layer and the outerlayer are made from different materials.
 5. The leading edge assembly ofclaim 1, wherein the outer layer and the inner layer are made from ametal, a ceramic material, or a ceramic matrix composite material. 6.The leading edge assembly of claim 1, wherein the outer layer and theinner layer have different densities.
 7. The leading edge assembly ofclaim 1, wherein at least two of the one or more compliant layers aremade from different materials.
 8. The leading edge assembly of claim 1,wherein at least one cooling passageway is defined through the innerlayer, the outer layer, and the one or more compliant layers.
 9. Theleading edge assembly of claim 8, further comprising: a coolant supplyin fluid communication with the at least one cooling passageway forselectively providing a flow of coolant through the at least one coolingpassageway.
 10. The leading edge assembly of claim 1, wherein theleading edge assembly is positioned on a wing, a nosecone, an enginecowl, an engine inlet, a fuselage, or a stabilizer of the hypersonicvehicle.
 11. A leading edge assembly for a hypersonic vehicle, theleading edge assembly comprising: an outer wall that is tapered at aforward end thereof; and a tip portion joined to the forward end of theouter wall, the tip portion comprising: a forward end that is curved, anaft end that is curved, and a plurality of structural layers and aplurality compliant layers alternately stacked with each other, whereinone of the plurality of structural layers is positioned at the aft endof the tip portion and conforms to the aft end of the tip portion,wherein one of the plurality of structural layers is positioned at theforward end of the tip portion and conforms to the forward end of thetip portion, wherein the plurality of compliant layers are curved andfacilitate movement between the plurality of structural layers, andwherein the plurality of structural layers each have a first coefficientof thermal expansion and the plurality of compliant layers each have asecond coefficient of thermal expansion, the second coefficient ofthermal expansion being greater than the first coefficient of thermalexpansion.
 12. The leading edge assembly of claim 11, wherein the outerwall includes a first wall section and a second wall section separatedby a chamber, and wherein the plurality of structural layers are curvedand extend between the first wall section and the second wall section.13. The leading edge assembly of claim 11, wherein the plurality ofstructural layers are made from different materials.
 14. The leadingedge assembly of claim 11, wherein the plurality of structural layersare made from a metal, a ceramic material, or a ceramic matrix compositematerial.
 15. The leading edge assembly of claim 11, wherein at leasttwo of the plurality of structural layers have different densities. 16.The leading edge assembly of claim 11, wherein at least two of theplurality of compliant layers are made from different materials.
 17. Theleading edge assembly of claim 11, wherein at least one coolingpassageway is defined through the plurality of structural layers and theplurality of compliant layers.
 18. The leading edge assembly of claim17, further comprising: a coolant supply in fluid communication with theat least one cooling passageway for selectively providing a flow ofcoolant through the at least one cooling passageway.
 19. The leadingedge assembly of claim 1, wherein the outer wall and the tip portionhave a uniform thickness.
 20. The leading edge assembly of claim 1,wherein both forward and aft ends of each of the one or more compliantlayers are curved.
 21. The leading edge assembly of claim 11, whereinthe outer wall and the tip portion have a uniform thickness.
 22. Theleading edge assembly of claim 11, wherein both forward and aft ends ofeach of the plurality compliant layers are curved.